A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine general includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, an airflow is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the compressor section and is then routed through the exhaust section, e.g., to atmosphere. In particular configurations, the turbine section is mechanically coupled to the compressor section by a shaft extending along an axial direction of the gas turbine engine.
The fan includes a plurality of blades having a radius larger than the core of the gas turbine engine. The fan and plurality of blades may also be mechanically coupled to the shaft such that they rotate along with the turbine. In certain configurations, the fan may be mechanically coupled to the shaft through a gear box, such that the fan can have a different rotational speed than the turbine shaft. A rotatable hub can be provided covering at least a portion of the fan and rotating along with the fan. Rotation of the plurality of blades generates thrust for the gas turbine engine and provides airflow to the compressor section of the core. Additionally, a plurality outlet guide vanes can direct an airflow from the blades to, e.g., reduce an amount of noise generated by the gas turbine engine and enhance performance of the gas turbine engine.
For at least some gas turbine engines, the fan is a variable pitch fan. It is desirable to vary the pitch of the fan blades by rotating each blade about respective pitch axes to further increase performance of the gas turbine engine. For example, a primary reason for changing blade pitch is to adjust the blade's angle of attack for optimal performance based on the present air speed of the aircraft and power level of the engine. In addition, the pitch of fan blades may be used to reverse the airflow through the core of the engine, thus providing reverse thrust to aerodynamically brake a landing aircraft.
In general, fan performance may be improved by increasing the number of blades. More specifically, it is desirable to maintain a blade solidity value of greater than one. Blade solidity is the ratio of the blade chord, represented by its length, over the blade pitch, which is the circumferential spacing of the blades at a given radius or diameter from the axial centerline axis. In other words, blade pitch is the circumferential length at a given diameter divided by the number of blades in the full fan blade row. Notably, when fan blades are designed with a solidity factor greater than one, adjacent blades will interfere with each other if they pass through flat pitch simultaneously.
Therefore, although fan blades with solidity greater than one are desirable for improved performance of the fan and engine, such a design can result in blade conflict when rotating into a reverse thrust configuration. Specifically, if all blades are rotated in unison, such that they rotate through flat pitch simultaneously, blade contact might occur.
Accordingly, a variable pitch fan for gas turbine engine including components allowing for asynchronous pitching is desirable. In addition, it is desirable that such a fan configuration and components remain lightweight and easy to assembly and service. More particularly, a fan for a gas turbine engine configured for asynchronous blade pitching while also allowing a higher fan blade solidity would be particularly useful.